Solar array augmented electrostatic discharge for spacecraft in geosynchronous earth orbit

ABSTRACT

Improved solar cell circuit layouts and cell structures that protect solar arrays located on spacecraft disposed in geosynchronous earth orbit from electrostatic discharge. An insulating material, such as RTV adhesive, for example, is used as a barrier that disposed in intercell gaps between solar cells. The use of the insulating material modifies sparking in the gaps caused by electrostatic discharge so that, while the spark still occurs, it has different non-destructive characteristics. The use of the insulating material causes no damage to other solar cell materials, such as a Kapton insulating layer or substrate used to support the solar cells. Furthermore, unique solar cell wiring schemes are provided that limit the voltage between adjacent solar cells to 50 volts or less.

BACKGROUND

The present invention relates generally to solar arrays used onspacecraft, and more particularly, to improved solar cell circuitlayouts and solar cell structures for protecting solar arrays located onspacecraft disposed in geosynchronous earth orbit from electrostaticdischarge.

During 1997, the assignee of the present invention launched fivehigh-powered spacecraft which generate over 10 kW of electrical power atthe beginning of their life. On two of those spacecraft there has beendamage to the solar arrays during the first year of operation. Extensiveanalysis and ground testing has demonstrated that a damage mechanismexists in which electrostatic discharges occurring between pieces ofcover glass and the solar cells on the solar arrays can be sustained bycurrent from the solar array itself. Depending on the physicalconstruction of the array, local heating can cause pyrolization of theinsulation which separates the solar cells from the conductivesubstrate, thus resulting in short circuits of individual strings ofsolar cells. Susceptibility to this phenomenon is likely to increasethroughout the industry as spacecraft power increases lead to largersolar arrays operating at higher voltages. However, analytical modelingand laboratory experimentation have verified the phenomenon andvalidated the preventative actions undertaken by the assignee of thepresent invention so that this phenomenon can be controlled on futurespacecraft.

It would therefore be desireable to provide for technical approachesthat protects solar cells on geosynchronously orbiting spacecraft fromdamage caused by electrostatic discharge. Accordingly, it is anobjective of the present invention to provide for improved solar cellcircuit layouts and cell structures for protecting solar arrays locatedon spacecraft disposed in geosynchronous earth orbit from electrostaticdischarge.

SUMMARY OF THE INVENTION

To accomplish the above and other objectives, the present inventioncomprises improved solar cell circuit layouts and cell structures thatprotect solar arrays located on spacecraft disposed in geosynchronousearth orbit from electrostatic discharge. The present invention providesfor the use of an insulating material as a barrier, such as using roomtemperature vulcanizable (RTV) adhesive or insulating material, forexample, disposed in intercell gaps between solar cells. The use of suchinsulating material modifies sparking in the gaps caused byelectrostatic discharge so that, while the spark still occurs, it hasdifferent non-destructive characteristics. The use of the insulatingmaterial causes no damage to other solar cell materials, such as Kaptoninsulating material disposed between the solar cells and the substrateto support the cells. Furthermore, unique solar cell wiring schemes areprovided that limit the voltage between adjacent cells to 50 volts orless.

In developing the present invention, a model was developed forspacecraft charging that shows that the solar panels of largegeosynchronous earth orbit communications satellites can exhibit an“inverse potential gradient” in which the solar cell cover glass chargesless negatively than the spacecraft body. The amount of inversepotential gradient is strongly dependent on the bulk resistivity of thecover glass.

A model of the arc discharge that can result from this potentialgradient was also developed that shows that a plasma created by thedischarge can trigger a sustaining arc, with current fed from the arrayitself. It has been found that there is a threshold cell-to-celldifferential voltage below which the sustaining arc cannot be created,which may be why it has not been a problem with spacecraft in the past.

The arc discharge model has been verified by testing at NASA LewisResearch Center. This verification has shown that arrays that have beenflown with high reliability for years (such as Intelsat VII, forexample) can fail if they are operated at sufficiently high cell-to-cellvoltage. Although heritage construction processes have been used, forboth high power 100 V GaAs and Si arrays developed by the assignee ofthe present invention, the threshold for damage has been shown to bejust at the limit of the normal operating range. The fact that the highpower Si arrays currently in use have not been damaged can be attributedto detailed construction differences and it may be strongly influencedby the use of cover glass having a relatively low bulk resistivity.

In implementing the present invention, a number of constructiontechniques have been developed to provide a margin against failure fromthe secondary arc. Once the failure mechanism was understood, combiningthese techniques provides a very large margin to prevent the arcdischarge phenomenon from occurring in the future. The corrective actionimplemented on future spacecraft to be launched by the assignee of thepresent invention will provide a safety margin significantly higher thanthat of previous solar arrays that have operated successfully for years.

Regarding the specifics in producing a solar array in accordance with apreferred embodiment of the present invention, the intercell voltagedifferential is lowered by 62.5% by rewiring the solar panels, thevoltage threshold at which damage can occur is increased by a factor of3 to 4 by adding the insulating material barrier between cells, and thecurrent available to the arc is decreased by a factor of 2 to 3 byadding solar cell string isolation diodes.

The susceptibility of solar arrays to electrostatic damage is a functionof their construction details, which determines how susceptible they areto damage. However, using the principles of the present invention, solararrays can safely be produced that operate at today's 10 kW powerlevels, and that will also operate at significantly higher power levelsthat will be used in the near future.

BRIEF DESCRIPTION OF THE DRAWINGS

The various features and advantages of the present invention may be morereadily understood with reference to the following detailed descriptiontaken in conjunction with the accompanying drawing, wherein likereference numerals designate like structural elements, and in which:

FIG. 1 illustrates a spacecraft anomaly that is corrected by the presentinvention;

FIG. 2 illustrates failure rate of a spacecraft as a function of timesince launch;

FIG. 3 illustrates spacecraft charging as a function of cover glassresistivity;

FIG. 4 illustrates cover glass differential charging as a function ofcover glass resistivity;

FIG. 5 illustrates cover glass resistivity;

FIG. 6 illustrates cover glass charging potential as a function ofdistance from the spacecraft main body;

FIG. 7 illustrates solar cells charge in response to space environment;

FIG. 8 illustrates that spacecraft charging causes a small arc to occurin the gap between the solar cells.

FIG. 9 illustrates the spacecraft arc charging triggers a sustaineddischarge driven by the array string current and voltage;

FIG. 10 illustrates a top view of the array showing an intercell gap;

FIG. 11 illustrates the use of an insulating barrier between cells inaccordance with the principles of the present invention;

FIG. 12 illustrates a test setup used to test the present invention;

FIG. 13 illustrates an example of solar array augmented dischargeleading to failure;

FIG. 14 illustrates a measured GaAs coupon failure threshold;

FIG. 15 illustrates a measured Si coupon failure threshold;

FIG. 16 illustrates a GaAs coupon failure threshold with the presentinsulating barrier installed;

FIG. 17 illustrates a Si coupon failure threshold with the presentinsulating barrier installed;

FIG. 18 illustrates a solar cell layout in accordance with the presentinvention wherein the cell-to-cell voltage is limited to 50 V or less;

FIG. 19 illustrates that the insulating barrier protects exposed Kaptoninsulator;

FIG. 20 illustrates that each solar cell string is isolated by diodes,limiting the current available to an arc;

FIG. 21 illustrates a exemplary solar cell circuit layout in accordancewith the present invention that is resistant to electrostatic discharge;and

FIG. 22 illustrates an exemplary solar cell circuit layout in accordancewith the present invention that provides for electrostatic dischargecurrent limiting.

DETAILED DESCRIPTION

Referring to the drawing figures, and in particular FIG. 1, anomaliesobserved on spacecraft 10 deployed by the assignee of the presentinvention are consistent with failures of strings 11 of solar cells 12on solar arrays 13 (solar panels 13). The symptoms of the failures arelow impedance shorts between solar cells 12 at different points within astring 11, and shorts between high voltage solar cells 12 and the arrayground. All of the anomalies occurred when other instrumented satellitesmeasured a charging environment characteristic of a solar substorm.

In a paper by Katz, et al. entitled “Mechanism for Spacecraft ChargingInitiated Destruction of Solar Arrays in GEO,” 36th AIAA AerospaceSciences Meeting and Exhibit, 1998, a theory and supporting laboratorydata were presented, showing how small, low energy, spacecraft chargingarcs on solar arrays can lead to larger, sustained discharges, in turnleading to permanent damage of the solar arrays 13. Since the Katz, etal. publication, additional work preformed by the assignee of thepresent invention has led to an understanding of the phenomenon, of theexistence of a cell-to-cell voltage threshold below which the sustainingarc cannot occur, and of the appropriate preventative action. Each stepof this theoretical understanding has been empirically verified.

Two of the spacecraft 10 developed by the assignee of the presentinvention use Gallium Arsenide (GaAs) solar cells 12. These are the twospacecraft 10 that have suffered damage, with 18 out of 80 circuits(22%) having been damaged. FIG. 2 shows the cumulative failure historyas a function of time since the launch of the spacecraft 10. As can beseen, the failure rate has decreased significantly with time.

The remaining three arrays 13 use Silicon (Si) solar cells 12. One ofthe 120 circuits on these spacecraft 10 has suffered an anomaly. Itssignature is different from the others described herein, and it is notpossible to tell from data available on the ground whether or not it isfrom the same cause. As will be shown, the likelihood of damage is afunction of many of the detailed construction parameters of the array13. Although the arrays 13 discussed herein are GaAs and Si, one shouldnot conclude that GaAs, per se, is any more susceptible to damage thanSi.

Spacecraft charging analysis using the NASA Charging Analyzer Program(NASCAP) is discussed in a paper by Katz, et al. entitled “TheCapabilities of the NASA Charging Analyzer Program,” Spacecraft ChargingTechnology—1978, NASA CP-2071, AFGL-TR-79-0082, edited by R. C. Finckeand C. P. Pike, p. 101, 1979. This paper identified the solar arrays 13as the probable site of spacecraft charging initiated arcs.

For decades it has been known that 30-50 keV electrons in themagnetosphere can cause electrostatic potentials of several thousandvolts to develop between different surfaces on the same spacecraft 10.Flight observations have shown that the large potential differences canlead to arc discharges. This phenomenon is discussed in papers byDeForest, entitled “Spacecraft Charging at Synchronous Orbit”, J.Geophys. Res., 77, p 651, 1972, Mullen et al., entitled “Scatha Surveyof High-Level Spacecraft Charging in Sunlight”, J. Geophys. Res., 91, p.1474, 1986, and Koons, entitled “Summary of Environmentally InducedElectrical Discharges on the P78-2 (SCATHA) Satellite”, J. Spacecraftand Rockets, 22, p 425, 1983. Modem geosynchronous communicationssatellites are typically covered with conducting surfaces to preventsurface potential differences and arcing. The only nonconductingsurfaces are usually pieces of solar cell cover glass 16 (shown in FIGS.7-11 and most clearly in FIG. 19).

The NASCAP program was used to analyze the electrical charging expectedon the spacecraft 10, and to locate possible discharge sites. Becausethe actual environment is unknown, the calculations were performed usingthe NASA recommended “worst case” charging environment discussed in apaper by Purvis et al. entitled “Design Guidelines for Assessing andControlling Spacecraft Charging”, NASA TP 2361, 1984. Parameters of the“worst case” charging environment are n_(e)=1.12×10⁶ m⁻³, T_(e)=12 keV,n_(i)=2.36×10⁵ m⁻³, and T_(I)=29.5 keV.

The analysis shows that a spacecraft 10 charges during substorms, evenwhen it is in sunlight. The reason is that photoemission from therelatively small conducting area that is sunlit is exceeded by thecharging currents to the large conducting areas that is in the dark ifthe front surface of the solar arrays 13 is steered towards the sun. Inthe presence of a charging environment, the chassis of the spacecraft 10charges negative at an initial rate of about −5 V/s, as is shown in FIG.3.

The front surface of the solar array 13, however, always faces the sun,and charge from the cover glass 16 is constantly bled off viaphotoemission (except during an eclipse, when the arrays 13 do notproduce power). This leads to an inverted potential gradient in whichthe pieces of the cover glass 16 are less negatively charged than theunderlying cell material and metallic interconnects. As is shown in FIG.4, the potential difference between the cover glass 16 and theunderlying cell increases at an initial rate of about 3 V/s. Thephotoemission current maintains the cover glass 16 near ambientpotential.

Initially, the resistivity of the cover glass 16 was modeled as beingeffectively infinite. Further examination, revealed however, that if thebulk resistivity, ρ, of the cover glass 16 was less than about 10¹³ Ω-m,then the maximum charge might be held below the threshold ofelectrostatic discharges. For example, for a cover glass 16 having aresistivity of 10¹³ Ω-m, the maximum voltage is limited as follows:$\begin{matrix}{V = {\left( {\rho \times t} \right)I}} \\{\left. {= {8 \times 10^{12} \times 125 \times 10^{- 6}}} \right)10^{- 6}} \\{{= {1000\quad V}},}\end{matrix}$

where ρ=weighted average of the resistivity (Ω-m) of the cover glass andadhesive, t=100 microns of glass and 25 microns of adhesive, and I=10⁻⁶A/m⁻².

FIG. 5 shows the bulk resistivity for various materials used in coverglass 16. As can be seen, the resistivity of CMZ cover glass 16, whichhas been used on GaAs arrays 13 produced by the assignee of the presentinvention, is high enough that it does not significantly limit thecharging voltage. The resistivity of 0213 cover glass 16, which is usedon both Si arrays and GaAs arrays now under construction by the assigneeof the present invention, is sufficiently low that it reduces themaximum charge potential. The exact value is a function of the chargingenvironment, as well as the exact temperature of the solar array 13 atthe time. This is one possible explanation for the fact that the Siarrays 13 have not experienced troubles on orbit. For future spacecraft10, by the assignee of the present invention is investigating switchingto CMX or other lower resistivity cover glass 16 as standard practice.

The final step of the charging analysis is the recognition that theelectrostatic potential is not constant along the array 13, increasingas the distance from the main body of the spacecraft 10 increases. FIG.6 illustrates this, for a snapshot two minutes after a charging event isinitiated, assuming infinite cover glass resistivity.

As can be seen in FIG. 6, the voltage on the fourth panel 13 is over 300V, decreasing to 140 V on the first panel 13. Of the 18 failures onorbit, 7 have occurred on the fourth panel 13, or outermost panel 13, 4on the third panel 13, and 2 on the first panel 13. The location of theremaining 5 failures is not known precisely, but it is known that theyare not on the first panel 13. The inference is clear. in that,discharges occur first on the outermost panel 13. Once a dischargeoccurs, the spacecraft 10 is neutralized and the charging isreinitialized. If discharging continues to occur preferentially on theoutermost panel 13, this will protect the inboard panels 13. The factthat failures occurred very early in life on the outermost panels 13 andhave decreased significantly in frequency suggests that the remainingportion of the solar arrays 13 may continue to operate normally for therest of the lifetime of the spacecraft 10.

The arcs described above have insufficient energy or currents to lead topermanent failures in the power system of the spacecraft 10. Furtheranalysis suggested that the short duration spacecraft charging arcscould trigger long duration discharges between solar cells 12; thesedischarges supported by the solar array current itself. The longduration discharges dissipate substantial energy, and can causepermanent damage to Kapton insulating material 19 (FIG. 19) on which thesolar cells 12 are mounted. The effect on the Kapton material 19 is toturn this high resistance polymer into a low resistance carbonized ash.This process, pyrolysis, was previously reported in a paper by Stueberet al., entitled “Evaluation of Kapton Pyrolysis, Arc Tracking, andFlashover on SiOx-Coated Polyimide, Insulated Samples of Flat FlexibleCurrent Carriers for SSF”, NASA-CR 191106, 1993. The net effect is toshort out the string 11 of solar cells 12, either between high and lowvoltage cells 12, or between cells 12 at different positions in thestring 11 and the underlying substrate 18 which is at spacecraft chassisground potential, as shown in FIG. 1.

A simple theoretical model was developed for generating a plasma by adischarge between the solar cell 12 and its cover glass 16. FIGS. 7-9sequentially illustrates what happens. First, the solar array 13 chargesdue to the space environment (FIG. 7). The insulated cover glass 16 isdischarged due to photoemission of electrons, leading to a differentialcharge between the cover glass 16 and the solar cells 12. Next, anelectrostatic discharge occurs between the cover glass 16 and a solarcell 12 (FIG. 8). This “triggering arc” 17 is known to occur at adifferential voltage of several hundred volts. The discharge containsvery little energy and is, of itself, harmless.

However, a plasma created by the triggering arc 17 can collect currentfrom the array 13 itself, leading to a “sustained arc” 17 a (FIG. 9).The energy in this discharge, the product of the discharge current andthe plasma voltage, can be substantial, and can be sufficient topyrolize the Kapton insulating material 19. Thus phenomenon has beencaptured on film in NASA Lewis Research Center (LeRC) tests describedbelow, and is quite spectacular.

The trigger current model is for spherical expansion of the plasmagenerated during solar cell discharge. The energy in the discharge, theproduct of the discharge current and the charging voltage, generates aplasma at the arc site by ionization. The plasma is assumed to expandspherically from the arc site at a velocity of 3×10⁴ m/s, a typicalvalue based upon other experimental studies. The electron and ion plasmadensities are shown to vary as 1/r², where r is the distance from thearc site.

The saturation current density collectible from the arc discharge plasmais estimated assuming that the electrons in the plasma have a Maxwellianvelocity distribution with a temperature of 1.5 eV, which is typical forlow discharge generated plasmas. The saturation current is the maximumcurrent that can flow into the plasma from the solar array 13 if thepotential of the collecting solar cell 12 is above the plasma potential.If the saturation current is greater than the maximum current in thesolar array 13, it will limit the current.

The maximum current carried by the trigger arc 17 is found byintegrating the plasma thermal electron current density over the exposedconduction surface area of adjacent cells 12, as shown in FIG. 10.

Maximum circuit current is found by integration over x, the distancealong the cell:$I_{\max} = {{h{\int_{- \infty}^{\infty}{\frac{j_{e}(1)}{g^{2} + x^{2}}{x}}}} = {\pi \frac{{hj}_{e}(1)}{g}}}$

where h is the cell height, g is the intercell gap distance, andj_(e)(1) is the current density at a distance of one meter from the arcsite.

For the geometry of the solar arrays 13 in orbit that were developed bythe assignee of the present invention, the saturation current was foundto be 2.6 A, for an assumed discharge voltage of 500 V. Coincidentally,the maximum current available from the solar array 13 is almostidentical (2.1 to 2.6 Amps per circuit for the GaAs arrays 13, and 2.3to 3.4 Amps per circuit for the Si arrays 13).

FIG. 11 illustrates the use of a protective insulating barrier 20 (RTVadhesive 20) installed in gaps 23 between cells 12 in accordance withthe principles of the present invention. Using the same modelingtechniques, and the geometry of GaAs arrays 13 developed by the assigneeof the present invention, the maximum current is reduced from 2.6 Ampsto 0.15 Amps, which is a reduction factor of 17.

Presented below is a summary of the sustaining current model developedby the assignee of the present invention.

Step 1: The plasma expands spherically from a spot (arc site) ofdiameter ≈10 μm, $n_{i} = {\frac{I_{i}}{{ev}_{i}4\pi \quad r^{2}}.}$

Step 2: The arc electron current flows through a hemisphere$j_{e} = {\frac{I_{e}}{2\pi \quad r^{2}}.}$

Step 3: The scattering is dominated by classical electron-ion collisions$\begin{matrix}{v_{ei} = {3.9 \times 10^{- 12}n_{e}\Lambda \quad T_{e}^{{- 3}/2}}} \\{\Lambda = {30 - {\ln \quad {\left( {n_{e}^{1/2}T_{e}^{{- 3}/2}} \right).}}}}\end{matrix}$

Step 4: The classical conductivity is given by σ=ε₀ω_(p)/ Vei.

Step 5: The voltage drop found by integrating over the radius${V(r)} = {\int_{r0}^{r}{\frac{I_{e}}{{\sigma 2}\quad \pi \quad r^{2}}{{r}.}}}$

The intercell voltage threshold effect will now be discussed. Thevoltage drop calculated in step 5 above is the potential drop in theplasma due to the ohmic resistance of the plasma to the dischargecurrent. The voltage drop is in the range of 40 V, for reasonableassumptions of emission spot radius and ion current. The solar array 13only collects electron current if the cell-to-cell bias voltage isgreater than the resistive drop in the arc plasma. The current collecteddrops exponentially for bias voltages less than the local plasmapotential.

Laboratory tests have been performed which validate this theory andwhich have led to a further understanding of the mechanisms that areinvolved. The tests successfully reproduced the failure symptomsobserved in orbit, and provided great insight into the details of thecontrolling mechanisms.

A simple model for the generation of a plasma by trigger arc dischargeof the cover glass 16 to the solar cell 12 has thus been developed. Thisanalytical model is supported by the experimental data from testingperformed at the NASA Lewis Research Center. In these tests, a solarpanel coupon is installed in a thermal vacuum chamber with the potentialbetween cells 12 adjustable externally.

The test setup, shown in FIG. 12, provides an inverted voltage gradientbetween the cover glass 16 and the solar cells 12 and their substrate18. A bias supply 25 is used to set the substrate and cell ground returnnegative relative to chamber ground. A plasma source 26 is used to floodthe cover glass 16 with a low energy (1-2 eV) low density plasma andthereby maintain the potential of the cover glass 16 near the chamberground. A solar array simulator (SAS) 27 is used to simulate the voltageand current from the solar array 13 and provides for the differentialcell voltages which are necessary to create the sustaining arc. The biassupply 25 charges a capacitor 28. The value of the capacitor 28 and thebias supply voltage determine the energy available to the arc discharge.

This arrangement is actually more representative of a low earth orbitenvironment, and is thought to represent a significantly worst casescenario for geosynchronous earth orbit. Failures that occurinfrequently in geosynchronous earth orbit can be systematicallyinduced, leading to rapid verification of protective measures. Thegeosynchronous earth orbit environment has also been simulated in asolar thermal bacuum chamber of the present assignee. However, arcing atvoltages below several thousand volts have not been induced. It appearsthat the phenomenon does not occur in a perfect vacuum, and a plasmamedium must be present to initiate the arc.

The amount of energy that can be stored during spacecraft charging isproportional to the capacitance of the solar cell 12 and cover glass 16combination. A capacitor 28 in the test setup shown in FIG. 12 is usedto control the energy of the trigger arc. Capacitance values up to twicethe capacitance of an entire solar panel 13 are used to demonstratemargin.

The number of arcs for a 30 minute period are plotted versus bias supplyvoltage. From these tests the arc threshold voltage between cover glass16 and solar cell 12 is determined. The test also measures the currentwaveforms for all arc discharges. The instrumentation allows the initialtrigger arc to be visible and the absence or presence of a sustainingarc can be observed. The SAS voltage is increased in steps to determinethe safe intercell voltage operating range.

Test results from the solar array 12 for a GaAs solar cell coupon withthe same construction as on the on-orbit spacecraft 10 is shown in FIG.13. More specifically, FIG. 13 illustrates solar array augmenteddischarge leading to failure. The solar cell bias was set to 80 V andthe solar array current limit set to 2.25 Amps, close to the limit ofthe actual on-orbit operating conditions. The current supplied by thesolar array 13 rises immediately to the current limit value. Thedissipation of this large initial circuit current at the collecting cell12 resulted in overheating and sustaining of the discharge beyond the100 μsec of the trigger arc. The result was a cell-to-cell andcell-to-substrate short.

A number of tests similar to the one described above have beenperformed. The results show that failures can be induced in both the Siand the GaAs arrays 13 of the type launched by the assignee of thepresent invention. As predicted by the model, there is a voltagethreshold value below which the trigger arc extinguishes itself withinμsec, although the exact value of the threshold cannot be determinedbecause it is dependent upon the precise geometry of the arc site.

FIGS. 14 and 15 show that the onset of failure is just at the boundaryof the on-orbit cell operating regime. No current to the solar cells 12was observed in any NASA LeRC testing for cell voltages below 60 V,which is consistent with the magnitude of the threshold predicted by themodel described above.

Testing of samples with the insulating (RTV) barrier 20 installed in thegap between cells 12, as shown in FIG. 1, has demonstrated theeffectiveness of the insulating (RTV) barrier 20, with the failurethreshold being increased significantly beyond the operating regime.This is shown in FIGS. 16 and 17. At the present time, both Si and GaAscoupons, manufactured in accordance with the principles of the presentinvention that will be used for future flights, are undergoing testingat NASA LeRC. The test procedure is set up to demonstrate margins ofsafety with respect to generation of a sustaining arc of at least afactor of 2 in voltage, and at least a factor of 2 in discharge energy(twice the capacitance equal to an entire panel 13). A minimum of 240arcs are recorded for each case in the test matrix.

Testing of a GaAs coupon with insulating (RTV) barrier 20 installed hasbeen completed. Discharge occurred at inverse gradient voltages of 290to 530 V, and discharge voltage increases with time which could indicatethat the threshold increases subsequent to arcs occurring at “weaker”locations.

The coupon experienced over 1000 arc discharges. The current recordingsduring the discharges showed that the insulating (RTV) barrier 20 issuccessful in limiting the flow of current into the plasma formed by thetrigger arc. Current from the solar array simulator was generally in the0.2 to 0.4 A range during the 50 to 100 μsec duration of the discharge.There were no cell-to-cell or cell-to-substrate shorts, and there wereno instances of sustained arcing being fed by the solar array simulator.The conclusion is that the insulating (RTV) barrier 20 is effective inpreventing the sustaining arc, as predicted by theory. In fact, previousexperimental samples subjected to testing at far higher voltages showedthat the ultimate damage occurred beyond the insulating (RTV) barrier 20and that much higher cell-to-cell voltages were required to initiatesuch damage.

For future high power spacecraft to be launched by the assignee of thepresent invention, three corrective actions have been undertaken, any ofwhich is sufficient to prevent damage from the phenomenon described inthis paper. Together, significant margin is demonstrated.

First, solar array panels 13 are wired in accordance with the principlesof the present invention so that the voltage between adjacent cells 12is 50 V or less, as is shown in FIG. 18. The previous high poweredspacecraft 10 had cases of 80 V (GaAs) and 75 V (Si) differentialsbetween adjacent cells. Extensive analysis is used to ensure that thisdifferential limit is not exceeded for various combinations of shuntedand unshunted strings 11 of solar cells 12, shadowing cases, and casesof failed strings 11. The assignee of the present invention typicallyoperates arrays at voltages up to 50 V, with excellent reliabilityresults.

Second, as shown in FIG. 11 and in detail in FIG. 19, the insulatingbarrier 20, such RTV adhesive 20 or insulating material 20, is insertedin all gaps 23 between cells 12, and for a distance of at least 10 mm inthe crossing gaps between series connected cells 12. As is shown in FIG.19, the structure of the improved solar array 13 comprises a substrate30 having an aluminum core 31 surrounded by a graphite skin 32. TheKapton insulating layer 19 is diposed on top of the substrate 30. The(RTV) insulating material 20 is disposed between the Kapton insulatinglayer 19 and bottom surfaces of the solar cells 12 and in the gaps 23between the solar cells 12. Cover glass 16 covers exposed surfaces ofthe the solar cells 12. Testing of sample solar arrays 13 during thedevelopment of this process has shown that it increases the thresholdfor damage to about 200 V.

Third, an as is shown in FIG. 20, for arrays 13 in orbit, there are asmany as five GaAs strings 11 in parallel, and as many as three Sistrings 11 in parallel, to form a “circuit”. Each circuit utilizes anindividual solar array drive assembly (SADA) slip ring, and is connectedto an individual switching shunt element (see FIG. 1). This arrangementprovides up to 3.4 Amps (Si) and 2.6 Amps (GaAs) to feed the discharge.

Future GaAs arrays will use larger cells 12, such that the individualstrings 11 will have current capability substantially identical to theSi arrays 13. Parallel strings 11 will be separated by diodes 22, as isshown in FIG. 20, on all unlaunched spacecraft 10 so that the mostcurrent available to an arc will be 1.1 A for both the Si and GaAspanels 13. This has the inherent benefit of a significant increase incalculated reliability, even without considering the phenomenondescribed in this paper, at a small penalty in efficiency of the solararray 13.

Additional details regarding the structures discussed with reference toFIGS. 18-20 are presented below. In particular, two high voltage solararray designs are shown in FIGS. 21 and 22, respectively. The presentinvention provides for techniques for designing solar panels 13 for useat high power levels. As was mentioned previously, prior solar paneldesigns that are to operated at voltages greater than 70 V, for example,had the potential to fail when the space environment causedelectrostatic discharge. The present invention allows use of solar panelvoltages over 70 volts. Solar panels 13 produced in accordance with theprinciples of the present invention are more robust to electrostaticdischarge.

A first exemplary circuit layout accordance with the present inventionis shown in FIG. 21 that is resistant to electrostatic discharge. Solarcell circuits to be used for spacecraft power systems may be assembledfrom all commonly used photovoltaic devices, including silicon, galliumarsenide and multi-bandgap cells. As is shown in FIG. 21, the solarcells 12 are arranged on the solar panels 13 in a pattern that preventsadjacent solar cells 12 from having voltage potential greater than 50volts. A spiral interconnection arrangement is used to seriallyinterconnect the solar cells 12 together. Furthermore, isolation orblocking diodes 22 are used to limit reverse current through a shortedcircuit.

Referring to FIG. 22, it illustrates a exemplary solar cell circuitlayout in accordance with the present invention that provides forelectrostatic discharge current limiting. When circuits are to beconnected in parallel, the prevention of reverse current through ashorted circuit should be limited to less than 1.5 Amps. The use of theisolation or blocking diodes 22 is one means for limiting the reversecurrent through the shorted circuit. The circuit layout shown in FIG. 22provides for electrostatic discharge current limiting. The addition ofinsulating material (the insulating barrier 20) disposed between thesolar cells 12, such as RTV insulating material 20 or adhesive 20, forexample, as is illustrated in FIG. 19, provides further protection fromelectrostatic discharge for voltages in excess of 100 volts.

Thus, improved solar cell circuit layouts including means for protectingsolar arrays located on spacecraft disposed in geosynchronous earthorbit from electrostatic discharge have been disclosed. It is to beunderstood that the above-described embodiments are merely illustrativeof some of the many specific embodiments that represent applications ofthe principles of the present invention. Clearly, numerous and otherarrangements can be readily devised by those skilled in the art withoutdeparting from the scope of the invention.

What is claimed is:
 1. A solar array comprising: a substrate; aplurality or groups of series-connected solar cells separated from eachother by gaps; an insulating barrier disposed between the substrate andthe solar cells, and disposed in the gaps between the groups ofscrics-connected solar cells, and disposed so as to project away fromthe gaps a predetermined distance into crossing gaps between seriesconnected solar cells of each group, and wherein the insulating barrieronly partially fills the crossing gaps; and cover glass disposed overthe plurality of solar cells to cover exposed surfaces thereof.
 2. Thesolar array of claim 1 wherein the substrate comprises an aluminum coresurrounded by a graphite skin.
 3. The solar array of claim 1 wherein theinsulating barrier comprises an insulating adhesive.
 4. The solar arrayof claim 1 wherein the insulating adhesive comprises room temperaturevulcanizable (RTV) adhesive.
 5. The solar array of claim 1 furthercomprising an insulating layer disposed between the substrate and theplurality of solar cells.
 6. The solar array of claim 1 wherein theinsulating layer comprises Kapton material.
 7. The solar array of claim1 wherein the predetermined distance is at least 10 mm in the crossinggaps between series cells.
 8. The solar array of claim 1 wherein thesolar cells are wired together so that the voltage between adjacentcells is 50 V or less.
 9. The solar array of claim 8 wherein the solarcells are wired together in a spiral interconnection pattern to seriallyinterconnect the solar cells together, and which prevents adjacent solarcells from having voltage potential greater than 50 volts.
 10. The solararray of claim 1 further comprising a plurality of isolation diodescoupled to respective outputs of interconnected ones of the solar cellsto limit reverse current through a shorted solar cell.
 11. A solar arraycomprising: a substrate; an insulating layer disposed on the substrate;a plurality of groups of series-connected solar cells separated fromeach other by gaps: an insulating barrier disposed between theinsulating layer and the solar cells, and disposed in the gaps betweenthe groups of series-connected solar cells, and disposed so as toproject away from the gaps a predetermined distance into crossing gapsbetween series connected solar cells of each group, and wherein theinsulating barrier only partially fills the crossing gaps; and coverglass disposed over the plurality of solar cells to cover exposedsurfaces thereof.
 12. The solar array of claim 11 wherein the substratecomprises an aluminum core surrounded by a graphite skin.
 13. The solararray of claim 11 wherein the insulating barrier comprises an insulatingadhesive.
 14. The solar array of claim 11 wherein the insulatingadhesive comprises room temperature vulcanizable (RTV) adhesive.
 15. Thesolar array of claim 11 wherein the insulating layer comprises Kaptonmaterial.
 16. The solar array of claim 11 wherein the predetermineddistance is at least 10 mm in the crossing gaps between series cells.17. The solar array of claim 11 wherein the solar cells are wiredtogether so that the voltage between adjacent cells is 50 V or less. 18.The solar array of claim 17 wherein the solar cells are wired togetherin a spiral interconnection pattern to serially interconnect the solarcells together, and which prevents adjacent solar cells from havingvoltage potential greater than 50 volts.
 19. The solar array of claim 11further comprising a plurality of isolation diodes coupled to respectiveoutputs of interconnected ones of the solar cells to limit reversecurrent through a shorted solar cell.